Method of selectively matching a turbine wheel and turbine nozzle assembly



May 24, 1966 A. H. BELL In METHOD OF SELECTIVELY MATCHING A TURBINEWHEEL AND TURBINE NOZZLE ASSEMBLY 4 Sheets-Sheet l Filed July 25, 19633,252,212 EEL A. H. BELL Ill May 24, 1966 METHOD OF SELECTIVELY MATCHINGA TURBINE WH AND TURBINE NOZZLE ASSEMBLY 4 Sheets-Sheet 2 Filed July 251963 www;

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May 24, 1966 A. H. BELL nl 3,252,212

METHOD OF' SELECTIVELY MATCHING A TURBINE WHEEL AND TURBINE NOZZLEASSEMBLY Filed July 25, 1965 4 Sheets-Sheet I5 A R IN VENTOR.

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May 24, 1966 H. BELL nl 3,252,212

METHOD OF' SELEC ELY MATCHING A TURBINE WHEEL AND TURBINE NOZZLEASSEMBLY Filed July 25, 1963 4 Sheets-Sheet 4 IN VEN TOR.

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United States Patent O M 3,252,212 METHOD F SELECTIVELY MATCHING A TUR-BINE WHEEL AND TURBINE NOZZLE ASSEM- BLY Albert H. Bell III, Birmingham,Mich., assignor to Chrysler Corporation, Highland Park, Mich., acorporation of Delaware Filed July 2S, 1963, Ser. No. 297,538 9 Claims.(Cl. 29-407) This invention relates generally to turbine engines andmore particularly to the fixed nozzle vanes and wheel blades employedtherein.

As Well known in the art, turbine engines are comprised Of a gasgenerator section consisting primarily of an air compressor, burnerportion and a compressor turbine wheel which obtains its energy from thegases fiowing out of the burner portion and drives the compressor. Someturbine engines, depending on their intended use, also include a powerturbine which when placed in the path of the flowing gases, downstreamof the compressor turbine, serves to provide additional work such as bydriving a propeller in a turbojet engine or driving some other outputshaft for land or water-based vehicles.

The ultimate output power developed by any turbine engine is primarilydependent on the output of the gas generator section which, in turn, isa function of the compressor speed of that particular gas generator.

In many instances it has been found that substantial variations inoutput horsepower exist as between any two turbine engines even thoughthe turbine engines are of the same design and rated output power. Suchvariations, which in some cases have been in the magnitude of thirtypercent (30%) power loss, exhibit themselves to the greatest extent inthe range of compressor speeds of seventy-five percent (75%) to ahundred percent (100%) of the designed or rated maximum compressorspeed. Further, these variations occur most frequently in turbineengines wherein the gas fiow through the fixed guide vanes or nozzlesand compressor turbine Wheel blades closely approaches sonic velocity asthe compressor speed approaches designed maximum speed.

As a consequence of such power variations, it becomes impossible topredict, with any reasonable degree of accuracy, the expected poweroutput of any particular turbine engine. Further, since such variationshave been found to exist between engines of identical design, it must beconcluded that the engine delivering the lesser power is, at least tosome degree, of lesser efficiency and therefore undesirable.

Now it has been discovered that if certain relationships between theelements comprising the 4turbine engine are considered as being criticaland such relationships are established during the process ofmanufacturing the turbine engine, that the resulting engines willproduce a reasonably predictable output horsepower which is consistentlyat a relatively higher and more eliicient value.

Accordingly, a general object of this invention is to provide, in aturbine engine, iixed guide vanes or nozzles and cooperating compressorturbine wheel blades having a relationship established therebetweenwhich enables the turbine engine to exhibit an overall higherefficiency.

Another object of this invention is to provide in a turbine engine,fixed guide vanes or nozzles and cooperating wheel baldes having arelationship therebetween which enables the turbine engine to achieve apredicted output horsepower which is within a reasonable range oftolerances.

Still another object of this invention is to provide a method forachieving a particular critical relationship between fixed guide Vanesor nozzles and cooperating tur- 3,252,212 Patented May 24, 1966 ICC binewheel blades of a turbine engine so as to obtain a turbine engine havinga predictable horsepower output. Another object of this invention is toprovide a method for achieving the above said critical relationshipwhich method includes the selective matching of a turbine wheel andcooperating turbine Wheel nozzle assembly.

A further object of this invention is to achieve the above said criticalrelationship by a method of selective matching of turbine Wheel andcooperating turbine wheel nozzle assembly which method includes matchingbased on the effective iioW areas of said turbine wheel and said turbinewheel nozzle assembly.

The invention as hereinafter disclosed in detail, contemplates a methodof selectively matching a turbine wheel to a cooperating turbine nozzleassembly which comprises the steps of creating a first iiow of asuitable fiuid through the nozzle assembly, determining from said flowof fiuid therethrough the effective flow area of said nozzle assembly,repeating the above steps with succeeding nozzle assemblies, classifyingsaid nozzle assemblies according to their respective equivalent flowareas, creating a second iiow of a suitable fluid through said turbineWheel, determining from said second flow of iiuid through said wheel theequivalent iiow area of said turbine wheel, repeating the steps withsucceeding turbine wheels, classifying each of the turbine wheelsaccording to their respective equivalent flow areas, and selectivelymatching a classified nozzle assembly to a classified turbine wheelaccording to a predetermined relationship of their respective equivalentflow areas.

Other objects and advantages of the invention will become apparent whenreference is made to the following description and drawings wherein:

FIGURE l is a cross-sectional view of a gas turbine engine adapted topropel driving wheels of a land based vehicle;

FIGURE 2 is a graph illustrating typical compressor curves obtained byplotting the ratio of compressor discharge pressure to compressor inletpressure against the weight-rate of air iiow through the compressor;

FIGURE 3 is a graph illustrating a typical output horsepower curve of aturbine engine;

FIGURE 4 is an enlarged fragmentary portion of the turbine engine ofFIGURE 1;

FIGURE 5 is a cross-sectional View taken generally on the plane of line5-5 of FIGURE 4, looking in the direction of the arrows;

FIGURE 6 is a cross-sectional view taken generally on the plane of line6 6 of FIGURE 5, looking in the direction of the arrows;

FIGURE 7 is a cross-sectional view taken generally on the plane of line7--7 of FIGURE 5;

FIGURE 8 illustrates a generally tubular conduit of convergingcross-sectional area;

FIGURE 9 is a graph illustrating the relationship between the change invelocity of iiiow through the conduit of FIGURE 8 for a correspondingchange in cross-sectional area; and

FIGURES l0 and ll illustrate, in cross-section, an arrangement fortesting nozzle assemblies and turbine wheels, respectively, inaccordance with the teachings ott this invention.

Referring now in greater det-ail t0 the drawings, FIG- URE l illustratesa turbine engine 10 adapted vlfor piropelling a land-based vehiclehaving driving wheels 12. The engine 10 is comprised of a housing 14having an air intake 16 and exhaust orifice 18. A combustion chamber 20,having any suitable fuel distribution means such las a fuel distributionring 22 therein, is located within the housing 14 between the compressor24 and compressor turbine wheel 26. Compressor 24 and iuzrbine wheel 26are connected to each other by means of a shaft 9 28 Iwhich may also beused for driving or in other ways operating selected engine accessoriessuch as the fuel control schematically illustrated at 30.

Compressor 24, as illustrated, is of the radial flow type and iscomprised of :a generally circular body portion 32 and cylindricalbearing portion 34 with a plurality of generally radially directedcompressor vanes 36 therebetween.

Turbine wheel 26 is provided with a plurality olf circumferentiallyspaced blades 38 extending radially outwardly from the wheel rim 40.Upstream of turbine wheel 26, the engine 10 is provided with a pluralityof radially directed guide vanes lor nozzles 42, sometimes referred toas stators, spaced oircumferentially about the body portion 44.

As compressor 24 is rotated, air is drawn in through inlet 16,compressed by vanes 36 and directed into the combustion chamber 20 wherefuel, supplied in accordance with a predetermined schedule, is burned soas to heat the gases therein. The gases flow from the burner chamber 20vthrough nozzles 42 which accelerate and caluse the gases to impinge uponthe turbine blades 38 in a manner causing rotative motion of turbinewheel 26r which, in turn, drives compressor 24.

The gases, continuing to flow towards exhaust orifice 18, pass between asecond set of guide vanes or nozzles 46, similar to nozzles 42, whichdirect the gases against blades 48 Iof an loutput power turbine wheel50. Turbine wheel S may be connected, a-s by any suitable powertransmission means 52, 54 to the driving wheels 12.

FIGURE 2 is a graph illustrating typical compressor curves 56, 58, 60,62, 64, 66, 68 and 70 obtained by plotting the ratio of compressordischarge pressure, P2, to,

inlet pressure, P1, against the air flow through the compressor inpounds per second. Each of the curves can be obtained by maintaining thecompressor at selected constant speeds While varying the pressure ratioP2/P1. The respective compressor speeds are indicated, for illustrativepurposes, near each curve in terms lof percentages of designed maximumcompressor speed.

The area above and to the left of dash line 72 represents that area inwhich compressor instability or surge is. encountered. Compressorinstability or surge refersy to a condition sometimes referred to ashunting, that is, those portions of the compressor curves which tend toflatten out and become more horizontal indicate that slight variationsin the ratio of P2/P1 result in comparatively large changes in the airflow which give rise to compressor instabil-ity. Curve '72 is determinedgenerally by connecting together the respective points, on theindividual compressor curves, at which compressor surge is firstencountered.

In addition to the compressor surge area, as discussed above, a furtherimportant consideration remains. That is, the rapidreduction :of thenumerical value of the ratio of P2/P1 as illustrated by the generallyvertically depending portions of each of the compressor curves 56through 70.A Since the compressor inlet pressure, P1, may be assumed tobe constant, the conclusion must be that the downstream or compressordischarge pressure, P2, decreases and the rate of .P2 decrease, asillustrated yby the slope of the depending portions lorf the compressorcurves, is very rapidas compared to an increment Iof change in the airflow. The reduction in comp-ressor discharge pressure, P2, is cau-sed bythe choking effect of the air within the compressor. Consequently, aline 74 drawn generally through the points of each Iof the compressorcurves wherein the rate of change of the ratio PZ/ P1 starts to becomerapid defines .an a-rea, generally -below and to the right of line 74,which miglht be referred to as the compressor choke area.

Accordingly, with turbine engines,-precaution1s are taken to assure theoperation of the compressor within the limits defined by the surge lorstall line 72 and choke line 74. In order to achieve this, it has beenthe practice,

generally, to aerodynamically match the compressor turbine wheel and itsassociated nozzle assembly, as a subassembly, to the compressor. This issometimes referred to as gross matching. In some instances other limitssimilar to the choke line 74 and the stall line 72 are determined andemployed for aerodynamic reasons. The aircraft industry has alsoemployed additional means for avoiding at least portions of thecompressor stal-l or surge areas by providing compressor bleed valveswhich at times and in accordance with selected operating parameters,vent some of the compressor air to the atmosphere.) A turbine enginehaving such a gross-matched compressor, lcompressor turbine wheel andnozzle assembly, so as to operate between the stall line 72 and chokeline 74, should exhibit a characteristic horsepower curve 76, obtainedgenerally Aby plotting output horsepower `and the percent of designedmaximum compressor speed along a logarithmic Y-axis and an arithmeticX-axis, respectively, of the graph of FIGURE 3.

However, it ha-s been found that substantial unpredictable variations inoutput horsepower still exist between gas turbine engines which havebeen constructed in accordance with the prior art.

v For example, with reference to FIGURE 3, i-t has been found that whileone turb-ine engine might deliver full rated power, another engine ofthe sarne design and rating might well develop substantially lessyhorsepower than that predicted. That is, a particular engine mightdeviate from the predicted mean curve 76, as at some point 80, and afterreaching a maximum output as at .point 82 (substantially lower than therated value represented by point 7S) steadily decrease to some point 84which may represent a developed horsepower less than that `of point 80.

Further, it V has been foundv that this phenomenon of horsepowervariation, or lost horsepower, will at times occur in the same gas`turbine engine which hadoriginally developed the full rated horsepower.stated, the phenomenon of unpredictable horsepower variation occurs mostfrequently in gas turbine engines wherein the velocity of gas flowI'through the compressor turbine and the compressor turbinenozzleassembly is'in the transonic rangeas the compresso-r approachesits designed maximum speed.

In the past, various proposals Ihave been tried inl unsuccessfulattempts to overcome theproblem of horse-- power variation. For example,the redesigning, modification and/or inspection of such areas aslubrication systems, gear trains, horsepower demands ofengine` drivenaccessories, alignment of turbinecomponents and engine air leakages inno Way alleviated; the problem ofy horsepower variation.

Therefore, assuming point-84 to;be the lowest horsepower ever achieved,it can be seen that turbine engines, constructed in accordance with Vtheprior art and having the compressor turbine and nozzle assembly grosslymatched, as a sulbassembly, to thecompressor, can be expected ltoproduce an output horsepower value ranging anywhere between the limitsdefined by points 78l and 84. In view of this, it becomes evident thatthe gross aerodynamic matching of the compressor turbine wheel and itsnozzle assembly, as a subassembly, to the compressor is in itselfinsufficient for assuring reasonable attainment of predicted horsepowerand that still another influencing factor, heretofore not apparent tothose skilledin the art,

remains as an important consideration in the design and` As waspreviouslyy sor turbine wheel 26. Preferably, the nozzle or stator vanes42 are made integrally with a main body p-or-tion 86 which is secured toa portion 44 of the general engine housing 14. This may be accomplishedas by means of a nut 88 internally threaded so as to cooperate with athreaded portion 90 in axially urging the stator body 86 against theradial surface 92 of a mounting shoulder 94. Radially outwardly, anannular shroud 96 is formed preferably integrally with the nozzle vanes42. The shroud 96, in cooperation with the outer surface 98 of thestator body 86, defines an annular passage 97 in which the varies 42 arelocated yfor the directional control ofthe gases flowing therethrough.

The 'compressor turbine Wheel 26, secured to shaft 28 for rotationtherewith, has its blades 38 generally Within the confines of shroud 96.The clearance between the outermost ends 100 of moving turbines blades38 and the inner stationary surface 102 of shroud 96 is kept to aminimum. The outer surface 40 of turbine wheel 26 and the inner surface102 of -shroud 96 also define an annular passage, generally coaxial withpassage 97, for the ow of gasses therethrough.

The flow area as between any two adjacent nozzle vanes 42 Will be of thesmallest cross-sectional area between such adjacent blades. For example,assuming that a plane indicated by line 6--6 of FIGURE 5 were passedthrough vanes 42 so as to have the cross-sectional area between thevanes a minimum area, the lflow area 103 therebetween could bedetermined, generally, from the dimension-s H1, W1 and B1 as illustrated-in FIGURE 6. Similarly, the ow area of the turbine blades 38 could bedetermined by passing a plane indicated by line 7 7 of FIGURE 5 throughpoints resulting in a minimum area 104. The flow area between suchturbine blades would then be determined, generally, by the dimensionsH2, W2 and B2 as illustrated in FIGURE 7. (Because of the relativelysmall clearances existing between the ends 100 of turbine blades 38 andthe inner surface 102 of stationary shroud 96, the height, H2, in FIGURE7 can -be considered as the distance fro-m surface 40 of turbine Wheel26 to the inner surface 102 of the shroud 96.)

As previously stated, the gross matching of the nozzle assembly 43 andcompressor turbine wheel, as a subassembly, to the compressor, eventhough well known -in the art, only `assures the engine of operationwithin the limits defined by the stall line 72 and the choke line 74 ofFIGURE 2. Such gross matchng does not, however, give any assurance thatthe `engine will attain the maximum horsepower as represented by poi-nt78 of FIGURE 3, but rather that the 'horsepower attained will be some-Where |between the desired point 78 and the low limit of point 84- Ithas been discovered Ithat what is still further required, in addition tothe gross matching with reference to the compressor, :is a critical orselective matching of the nozzle assembly 43 to the compressor turbinewheel assembly 26 based on the respective effective iiow areas of each.

' Referring to FIGURES 4, 5 and 6 it can be seen that if plane 6-6determines the smallest cross-sectional area between two adjacent nozzlevanes 42 then the cross-sectional areas taken at local points upstreamof the plane 6 6 will define progressively larger areas generally as thedistance between the local area under consideration and the minimum4area 103 increases. In other words, the progressively larger areas canbe considered as defining, generally, the converging conduit 120 ofFIGURE 8l wherein the throat area, AT, would be a dimensional equivalentof the minimum area 103 defined by W1, B1 and H1 of FIGURE 6 and thevarious local areas, Am, AL2, ALS, etc., would constitute the increasingcross-sectional areas upstream of the minimum area 103. Further, just asthe throat area, AT, will determine the maximum volume rate of air fiowthrough conduit 120, so will the minimum area 103 between nozzle vanes42 determine the maximum volume rate of gas iiow therethrough for anygiven condition. Generally, it can be said that the absolute maximumrates in both arrangement will occur when the velocity of lair owthrough the most restricted areas is sonic.

Assuming that conduit is under a condition of sonic flow, that is, thevelocity of air fiow through throat 122 is Mach-one, a relationship canbe established -between the respective local areas and the velocity offlow through such areas. For example, an exponential curve, 124illustrated in FIGURE 9 can be obtained by graphically plotting theratio of the local cross-sectional areas, AL, to the throat area AT,against the velocity of flow at that particular local area.

From an inspection of curve 124, it can be seen that when the ratio ofareas, AL/AT, is 1.000 then the velocity of the flow Iat the local 4areais equal to Mach-one. It should also be observed, however, that thechange in local air velocity, Mach No., is large as -compared to arelatively small change in the local cross-sectional area, as refiectedby AR, in that range Where the ratio of areas, AL/AT, approaches 1.000.

Even though curve 124 of FIGURE 9 has been discussed primarily withreference to the converging conduit 120 of FIGURE 8, the principlesinvolved apply equally well to the nozzle assembly 43. That is, -aspreviously stated, the minimum area 103 determined by W1, B1 and H1 isthe equivalent `of the throat area, AT, of throat 122, and the variouslocal cross-Sectional areas upstream of the minimum nozzle Iarea 103vare the equivalents of the local cross-sectional are-as Am, AL2, ALs,etc.

Accordingly, it becomes evident that if the size of the minimum nozzlearea 103 is calculated to transmit gas therethrough at sonic velocitiesand at Ia specific weightrate of flow, that very slight deviations fromthe calculated size `of the minimum nozzle yarea as illustrated byFIGURE 9 will result in relatively large reductions in the velocity ofair or gas fiow through the minimum nozzle area 103. The immediateeffect of such velocity losses is a drastic reduction in the energyavailable for driving the compressor turbine Wheel 26 and ultimately thepower turbine 50.

It has been discovered that in some instances turbine engines, having fanozzle assembly wherein the blade surfaces 108 and 110 (FIGURE 5) were.003 inch further away from each other than the ideal calculateddimension, experienced about a ten percent (10%) power loss. It has alsobeen -found that such turbine engines could be made to produce theirfull rated horsepower if the nozzle assembly and compressor turbinewheel 26 in the engine ywere replaced by a nozzle assembly andcompressor turbine wheel which were selectively matched to each other.

For example, let it be assumed that it has been empirically determinedthat a particular engine requires the effective fiow area of the nozzleassembly and the effective iiow :area of the compressor turbine wheel tobe in-an ideal ratio of 1.000`to 1.500, respectively, as determined at aparticular pressure differential across each. (The effective orequivalent flow areas are determined by total flow and velocityexperienced through the entire nozzle assembly 43, that is the summationof the effective fiow areas between the individual nozzle vanes, or theentire compressor turbine wheel, as `the case may be.)

It is well known that dimensional tolerances are required in anymanufacturing operation. Therefore, it becomes impossible to randomlychoose any one nozzle yassembly out of many and similarly choose acompressor turbine wheel and have lany degree of assurance that the twowill have their respective equivalent flow areas in precisely theassumed ideal ratio of 1.000; 1.500. Therefore, the very tolerancesnecessary for the production of the nozzle assembly and compressorturbine wheel cause the respective equivalent flow areas to vary to thedegree sufficient to cause ow velocity losses as described withreference to FIGURES 4-9. Such variations resulting from themanufacturing tolerances also exhibit themselves in causing therespective equivalent ow areas to vary from the assumed ideal ratio of1.000; 1.500.

However, the discovery of the cause of .the problem of variations in`and loss of engine power, does not in and of itself suggest anecessarily practical solution of that problem because dimension-almanufacturing tolerances cannot be eliminated nor can they be reduced toa degree which results in prohibitive manufacturing costs.

It has been discovered that the solution of the problem resides in theselective matching of a nozzle assembly to a cooperating compressorturbine wheel by means of matching the ow performance of each of them.

For purposes of illustration, let the following be assumed to be theideal constants:

(l) Ratio of the total nozzle assembly ow area to thev ytotal compressorturbine wheel ow area=1.000/ 1.500;

(2) Total nozzle assembly ilow area=10.000 sq. inches;

(3) Total compressor turbine wheel flow area: 15.000 sq.

inches;

(4) Velocity of gas ow through nozzle yassembly at full rated enginepower=Mach 0.94;

(5 Velocity of gas ow through compressor turbine wheel at full ratedengine power=Mach one.

Further, let it be assumed that because of the small dimensionaldifferences obtained due to reasonable manufacturing tolerances, thatrelatively large changes occur (as previously discussed with referenceto FIGURES 8 and 9) in the equivalent flow areas of both the nozzleassembly and the compressor turbine, so as to result in the following:

(6) The effective or equivalent total nozzle assembly ow area varies, asbetween assemblies, from 9.700 to 10.300 sq. inches and (7) Theeffective or equivalent total compressor turbine ow area varies, asbetween turbines from 14.500 to 15.500 sq. inches.

With the above assumptions it can be seen that a nozzle assembly andcompressor turbine wheel randomly selected from a large group of eachcould yield the following equivalent flow area combinations having thecorresponding indicated equivalent flow area ratios, which would stillsatisfy the gross mat-ching requirements asA between the compressor andthe nozzle assembly and compressor turbine as a subassembly:

Combinations I and IV result in equivalent 110W areas which are verynear the assumed ideal area ratios.

It has been discovered that, in gas turbine engines wherein the velocityof motive gas ow through the nozzle .assembly or through the compressorturbine is in the range of Mach 0.9 to Mach 1.() as the compressorapproaches its designed maximum speed, the numerical equivalent of theactual area ratio may Vary in the range of plus or minus 1.0% from thenumerical equivalent of the ideal area ratio. Therefore, combinations Iand IV would result in the engine devoloping its full rated horsepowerat point 78 of the Ihorsepower curve of FIGURE 3.

However, the same nozzle assemblies combined in re- Verse order with thecompressor turbine wheels, as illustrated by combinations II and III,result in equivalent flow area ratios having a substantial deviationfrom the assumed ideal equivalent flow area ratio. Combinations II andIII, even though satisfying the gross matching requirements of thecompressor, would, nevertheless, cause the engine to produce an outputhorsepowervmuch less than the rated output of the engine such asrepresented by point 84 of the curve of FIGURE 3.

Accordingly, in order to insure proper engine perform- Iance so as toconsistently achieve a reasonably predictable horsepower output itbecomes necessary to selectively match the nozzle .assembly andcompressor turbine to each other.

A method of selectively or critically matching the nozzle assembly andcompressor turbine is disclosed with reference to the apparatus ofFIGURE 10 which is comprised of an outer housing 126 provided with acover 128 at one end thereof and 'an inlet passage 130 at the other end.Conduit 132, having a valve 134 serially connected therewith,communicates at its one end with a source 135 of pressurized air whileits other end serves to retain an orice plate 136 in proper position soas to maintain the orifice 138 within the path of the air flowing fromconduit 132 and into the plenum chamber 140 defined generally by housing126 and cover 128.

Conduits 142 and 144 are provided so as to communicate the pressures onopposite sides of the orifice plate 136 to a differential pressure gage146. Conduits 142, 144, orifice 138 and pressure gage 146 provide meansfor determing the mass rate of flow through the orice 138 and thereforethe entire system. The mass rate of W, W, can be determined by theapplication of the following general compressible mass flow equation:

Where:

(1) G=flow per unit of area;

(2) W=mass rate :of flow;

(3) A=Crosssectional flow :area of orice 138;

(4) C=the discharge coefficient of orifice 138 (a calibrated function ofthe Reynolds number as well known in the art;)

(5) P0=P2=total pressure upstream of orice 138;

(6) M=molecular weight of the fluid flowing;

(7) Tu=total temperature upstream of orifice 133;

(8) g=acceleration due to gravity;

(9) n.=the ratio of the specific heat of the uid at constant pressure,to the specific heat of the uid at constant volume;

(10) =universal gas constant; and

(11) 11=pressure ratio of downstream static pressure,

`Ps2 to upstream total pressure, Po A testing fixture 152 having anannular passage 154 formed therein is :secured insealing engagement withcover 128. A plurality of guide vanes 156 are provided within theannular passage 154 so as to impart a swirling motion to the air passingtherethrough in order to simulate the langle ofair impingementexperienced by the nozzle vanes in the actual engine. The nozzleassembly 43 to be tested is mounted, as illustrated at the exit of theannular passage 154.

During testing ofthe nozzle assembly 43 valve 134 is opened to thedegree necessary to establish a desired differential, AP, between theplenum chamber pressure, P3, and the barometric pressure, PB. Thispressure differential and effective cross-sectional ow area of thenozzle assembly will then determine the velocity of ow through thenozzle assembly.

Since nozzle assembly 43 and orifice 138 are in series withv each other,it becomes evident that the mass-rate of fiow, W, must be the same forboth the nozzle assembly 43 and the orifice 138. Accordingly, once themass-rate of flow, W, is determined for orifice 138, the same massrateof flow can be employed in the general compressible mass ow equation asapplied to the nozzle assembly. The terms of the equation, as applied tothe nozzle assembly will have the following meanings:

(l) A=actual cross-sectional flow area of nozzle assembly 43;

(2) Wzmass-rate of flow through the nozzle assembly (the same as themass-rate of flow through orifice 138);

(3) P0=P3=pressure upstream of the nozzle assembly;

(4) m=molecular weight of the fluid;

(5) T0=total temperature upstream of the nozzle assembly;

(6) C=the discharge coefficient of the nozzle assembly;

(7) g=acceleration due to gravity;

(8) n,=the ratio of the specific heat of the gas at constant pressure,Cp, to the specific heat of the gas at constant volume, Cv;

(9) R=universal gas constant; and (l0) r=pressure ratio of downstreampressure, PB, to

upstream pressure, P3.

In applying the general equation to the nozzle assembly, the equationmay be rearranged into the following form:

It should also be noted that W, m, To, g, n and R are the sarne valuesas when the general equation was applied to the orifice 138. Further, r,is determined from the pressure differential, Pg-PB. The only valueswhich are not known are C and A. However, the solution of the equation,as rearranged above, yields a value which is the product of C, thedischarge coefficient, and A, the actual cross-sectional flow area or,in -other words, the effective or equivalent flow area of the nozzleassembly 43 which is precisely the value desired.

Similarly, the compressor turbine wheel 26 can be secured at the outletend of a test fixture 158 (FIGURE 1l) which also has an annulus 154 andguide vanes 156. A shroud 162 located about turbine wheel 26 and securedatop fixture 158 is provided in order to duplicate the confining effectthat the shroud 96 will have within the engine. The remaining portion ofthe apparatus of FIG- URE ll can be made so as to be identical with thatdisclosed in FIGURE 10.

The general procedure employed in testing the nozzle assembly isfollowed in the testing of the turbine wheel 26. That is, the valve 134is opened to the degree necessary to again establish the differential,AP, between P3 and PB.

The mass-rate of flow, W, through orice 138 is then determined by theapplication of the general compressible mass ow equation. Once, W, isdetermined it then becomes possible to solve the same equation, asapplied to the compressor turbine wheel, for the combined term, CA, (asdiscussed with reference to the nozzle assembly of FIGURE 10) of theturbine wheel 26.

By employing the above method it then becomes possible to test aplurality of nozzle assemblies, one at a time, and determine and recordthereon the equivalent or eective flow area exhibited by that particularnozzle assembly. Likewise, the compressor turbine wheels can be testedand the equivalent flow area, so determined, recorded on the individualcompressor turbine wheels. It then becomes .a matter of simplyselectively matching a nozzle assembly and compressor turbine wheel,based on their respective equivalent ow areas, which will result in anequivalent flow area ratio within the critical limits defined on bothsides of the ideal equivalentqow area ratio. Such a selectively matchedset, when placed within the turbine engine will insure the engine of notonly operating within the limits of the choke and stall lines of FIGURE2 as achieved by gross aerodynamic matching to the compressor, but willfurther insure the engine of producing an output horsepower at leastvery closely approaching the designed maximum output horsepower asrepresented by point 78 of FIGURE 3.

Further, it has been discovered that as between any two nozzleassemblies `or compressor turbine wheels, slight differences may existbetween the respective curves defining the mass-rate of gas flowtherethrough as a function of the pressure differential across suchnozzle assemblies or compressor turbine wheels. This is believed due tothe slightly varying aerodynamic characteristics of the particularnozzle .assembly or compressor turbine wheel under consideration.Accordingly, the method for determining and critically selecting aparticular nozzle assembly to a cooperating compressor turbine wheel canbe modified so as to even avoid such slight discrepancies arising out ofsuch varying aerodynamic characteristics.

That is, preferably, when the ideal equivalent flow areas areempirically determined, they should be determined by employing apressure differential substantially equivalent to that pressuredifferential which the nozzle assembly or the compressor turbine wheel,as the case may be, will experience within the engine as during fullrated power operation or maximum designed compressor speed.

Subsequently, the pressure differential employed during calibration ofnozzle assemblies would vary from that pressure differential employedfor Calibrating compressor turbine wheels. For example, with nozzleassemblies, a sufficient pressure differential may be indicated whenpressure gage indicates a value which is 0.75 times the then existingbarometric pressure, whereas, in the case of the compressor turbineWheel a reading of 1.17 times the then existing barometric pressure mayindicate a proper pressure differential.

In each event, however, the calibration is conducted employing apressure differential substantially equivalent to the pressuredifferential actually experienced by the test piece under conditions ofactual engine operation thereby avoiding any inuencing factors whichmight otherwise exhibit themselves due to the aerodynamiccharacteristics of the turbine wheel or nozzle assembly.

The drawings and the foregoing specification constitute a description ofthe invention in such terms as to enable any persons skilled in the artto practice the invention, the scope of which is indicated by theyappended claims.

I claim:

1. A method of selectively matching a gas turbine engine compressorturbine wheel to the compressor turbine wheel nozzle assembly,comprising the steps of creating a first pressure differential acrosssaid nozzle assembly substantially equivalent to a predeterminedpressure differential which said nozzle assembly will experience withinsaid engine during selected periods of normal engine operation,employing said first pressure differential to flow a suitablegasvthrough said nozzle assembly, determining from said pressuredifferential and the Volume rate of flow of said gas therethrough theeffective flow area of said nozzle assembly, repeating the above stepswith succeeding nozzle assemblies, classifying each of the nozzleassemblies according to their respective effective flow areas, creatinga second pressure differential across said turbine wheel substantiallyequivalent to a second predetermined pressure differential which saidturbine wheel will experience within said engine during selected periodsof normal engine operation, employing said second pressure differentialto ow a suitable gas through said turbine wheel, determining from saidsecond pressure differential and the volume rate of ow of said gasthrough said turbine wheel the effective flow area of said turbinewheel, repeating the steps with succeeding turbine wheels, classifyingeach of the turbine ll l Wheels according to their respective effectiveflow areas, and selectively matching a nozzle assembly to turbine Wheelso as to have the numerical value of the ratio of their respectiveeffective flow areas within at least one percent of the numerical valueof the ratio of their respective ideal flow areas.

2. A method of selectively matching a gas turbine engine lcompressorturbine wheel to the compressor turbine wheel nozzle assembly,comprising the steps o f creating a first pressure differential acrosssaid nozzle assembly substantially equivalent to a predeterminedpressure differential which said nozzle assembly will experience Withinsaid engine during selected periods of normal engine operation,employing said first pressure differential to flow a suitable gasthrough said nozzle assembly, determining from said pressuredifferential and the rate of fiow of said gas therethrough the effectiveflow area of said nozzle assembly, repeating the above steps withsucceeding nozzle assemblies, classifying said nozzle assembliesaccording to their respective equivalent flow areas, creating a secondpressure differential across said turbine wheel substantially equivalentto a second predetermined pressure differential which said turbine wheelwill experience Within said engine during selected periods of normalengine operation, employing said second pressure differential to flow asuitable gas through said turbine Wheel, determining from said secondpressure differential and the rate of flow of said gas through saidturbine wheel the equivalent flow area of said turbine Wheel, repeatingthe steps with succeeding turbine Wheels, classifying each of theturbine Wheels according to their respective equivalent flow areas, andselectively matching a classified n-ozzle assembly to a classifiedturbine wheel so -as to have the numerical value of the ratio of theirrespective equivalent flow areas Within a predetermined range of thenumeral value of the ideal ratio of their respective ideal flow areas.

3. A method of matching a gas turbine engine turbine wheel to acooperating turbine Wheel nozzle assembly each of which is comprised ofa plurality of circum- -ferentially spaced radially directed vanesdefining respective gas flow areas, comprising the steps offsecuringsaid nozzle assembly to an outlet of a relatively large plenum chamber,directing a fiow of suitable pressurized gas to an inlet of said plenumchamber, throttling the flow of said suitable pressurized gas to thedegree necessary to establish ya first pressure differential between theinterior of said plenum chamber and the ambient pres,-`

sure which is substantially equivalent to a predetermined pressuredifferential which said nozzle assmbly will experienec Within saidengine vduring selected periods of normal engine operation, measuringthe volume rate of gas flow supplied to said plenuml chamber,determining from said first pressure differential and .said volume rateof gas flow the equivalent ow area of the actual nozzle assembly gasflow area, terminating the flow of pressurized gas to said plenumchamber, removing said nozzle assembly from said outlet, classifyingsaid nozzle as-` sembly according to its equivalent flow area, repeating the `above steps with succeeding nozzle assemblies, secur-v ing aturbine wheel to an outlet of a relatively large plenum chamber,directing a fiow of suitable pressurized gas to an inlet of said lastmentioned plenum chamber, throttling the flow of the last mentionedpressurized gas to the degree necessary to establish a second pressuredifferential between the interior of said last mentioned plenum chamberand the ambient pressure which is substantially equivalent to `apredetermined pressure differ-v ential which said turbine Wheel Willexperience Within said engine during selected periods of normal engineoperation, measuring the volume rate of flow of said last mentionedpressurized gas supplied to said last mentioned plenum chamber,determining from said second pressure` differential and said lastmentioned volumey rate of gasA flow the equivalent area of the actualturbine Wheel gas flow area, terminating the flow of said last mentionedpressurized gas to said last mentioned plenum chamber, removing saidturbine Wheel from said last mentioned outlet, classifying said turbinewheel according to its equivalent flow area, repeating the steps withsucceeding turbine Wheels, and selectively matching a classified noz-Zle assembly to a classified turbine wheel according to a predeterminedrelationship of their respective equivalent flow area classifications.

4. A method of selectively matching a gas turbine engine compressorturbine Wheel to the compressor turbine wheel` nozzle assembly,comprising the steps of creating a first pressure differential acrosssaid nozzle assembly substantially equivalent to a predeterminedpressure differential which said nozzle assembly will eX- perienceWithin said engine during selected periods of normal engine operation,employing said first pressure differential to flow a suitable gasthrough said nozzle assembly, determining from said` pressuredifferential and the rate of flow of said gas therethrough the effectivefiow area of said nozzle assembly, repeating the above steps Withsucceeding nozzle assemblies, classifying said nozzle assembliesaccording to their respective equivalent fiow areas, creating a secondpressure differential across said turbine wheel substantially equivalentto a second predetermined pressure differential which said turbine Wheelwill experience Within said engine during selected peri-ods of normalengine operation, employing said second pressure differential to flow asuitable gas through said turbine wheel, determining from said secondpressure differential and the rate of flow of said gas through saidturbine Wheel-the equivalent flow area of said turbine WheeLrepeatingthe steps With succeedingturbine Wheels, classifying each of the turbineWheels according to their respective equivalent flow areas, andselectively matching a classified nozzle assembly to a classifiedturbine wheelraccordingto a predetermined relationship of theirrespective equivalent flow areas. t

V5. A method lof fselectively matching a gas turbine engine compressorturbine Wheelto the compressor turbine wheel nozzle assembly, comprisinglthe steps of creating a first pressure differential across said nozzleassembly substantially equivalent to a predetermined pressuredifferential which said nozzle assembly will experience Within saidengine during designed maximum engine speed operation, employing saidfirst pressure differential to flow a suitablegas through said nozzleassembly, determiningfrorn said pressure differential and the rate offlow of said gas therethrough the effective flow area kof said nozzleassembly,repeating the above steps with succeeding nozzle assemblies,classifying said nozzle assemblies according to their respectiveequivalent iiow areas, creating a second pressure differential acrosssaid turbine Wheel substantially equivalent to a second predeterminedpressure, differential which said turbine Wheel will experience withinsaid engine during designed maximum engine speed operation, employingsaid second pressure differential to flow a suitable gas through saidturbine Wheel, determining fromV said second pressure differential andthe rate of flow of said gas through said turbine Wheel ythe equivalentiiow area of said turbine wheel, repeating they steps with succeedingturbine wheels, classifying each of the turbine wheels according totheir ,respective equivalent flow areas, and selectively matching a;classifiednozzle Aassembly to a classified turbine wheel according to apredetermined relationship of their respective equivalent flow areas.

o 6. A method of selectively-matching a gas turbine engine compressorturbinel Wheel to the compressor turbine Wheel nozzle assembly,comprising the steps of creating a first pressure differential acrosssaid nozzleassembly, employing said first pressure differential to flowa suitable gas through said nozzle assembly, determining from said firstpressure differential .and the rate of flow of said gas therethrough theeffective flow area of said nozzle assembly, repeating thev above stepswith succeeding nozzle assemblies, classifying said nozzle assembliesaccording to their respective equivalent flow areas, creating a secondpressure differential across said turbine wheel substantially equivalentto said first pressure differential, employing said second pressuredifferential to flow a suitable gas through said turbine wheel,determining from said second pressure differential and the rate of flowof said gas through said turbine Wheel the equivalent flow area of saidturbine Wheel, repeating the steps With succeeding turbine wheels,classifying each of the turbine Wheels according to their respectiveequivalent flow areas, and selectively matching a classified nozzleassembly to a classified turbine wheel according to a -predeterminedrelationship of their respective equivalent flow areas.

7. A method of selectively matching a gas turbine engine compressorturbine Wheel to the compressor turbine wheel nozzle assembly,comprising the steps of creating a first pressure differential acrosssaid nozzle assembly, employing said first pressure differential to flowa suitable fluid through said nozzle assembly, determining from saidfirst pressure differential and the rate of flow of said fluidtherethrough the effective flow area of said nozzle assembly, repeatingthe above steps with succeeding nozzle assembles, classifying saidnozzle assemblies according to their respective equivalent flow areas,creating a second pressure differential across said turbine wheelsubstantially equivalent to said first pressure differential, employingsaid second pressure differential to flow a suitable fluid through saidturbine wheel, determining from said second pressure differential andthe rate of flow of said fluid through said turbine wheel the equivalentflow area of said turbine wheel, repeating the steps with succeedingturbine wheels, classifying each of the turbine Wheels according totheir respective equivalent flow areas, and selectively matching aclassified nozzle assembly to a classified turbine wheel according to apredetermined relationship of their respective equivalent flow areas.

8. A method of selectively matching a gas turbine engine compressorturbine wheel to the compressor turbine Wheel nozzle assembly,comprising the steps of creating a first pressure differential acrosssaid nozzle assembly substantially equivalent to a predeterminedpressure differenttial which said nozzle assembly will experience Withinsaid engine during selected periods of no-rmal engine operation,employing said first pressure differential to flow a suitable fluidthrough said nozzle assembly, determining from said pressuredifferential and the rate of flow of said fluid therethrough theeffective flow area of said nozzle assembly, repeating the above stepswith succeeding nozzle assemblies, classifying said nozzle assembliesaccording to their respective equivalent flow areas, creating a secondpressure differential across said turbine Wheel substantially equivalentto a second predetermined pressure differential which said turbine wheelwill experience Within said engine during selected periods of normalengine operation, employing said second pressure differential to flow asuitable fluid through said turbine wheel, determining from said secondpressure differential and the rate of flow of said fluid through saidturbine Wheel the equivalent flow area lof said turbine wheel, repeatingthe steps with succeeding turbine wheels, classifying each of theturbine wheels according to their respective equivalent flow areas, andselectively matching a classified nozzle assembly to a classifiedturbine wheel according to a :predetermined Arelationship of theirrespective equivalent flow areas.

9. A method of selectively matching a turbine Wheel to a cooperatingturbine nozzle assembly, compris-ing the steps of creating a first flowof a suitable fluid through said nozzle assembly, determining from saidflow of fluid therethrough the effective flow area of said nozzleassembly, repeating the above steps with succeeding nozzle assemblies,classifying said nozzle assemblies according to their respectiveequivalent flow areas, creating a second flow of a suitable fluidthrough said turbine wheel, determining from said second flow of fluidthrough said Wheel the equivalent flow area of said turbine wheel,repeating the steps with su-cceeding turbine Wheels, classifying each ofthe turbine wheels according to ltheir respective equivalent flow areas,and selectively matching a classified nozzle assembly to a classifiedturbine wheel according to a predetermined relationship of theirrespective equivalent flow areas.

References Cited by the Examiner UNITED STATES PATENTS 2,796,658 6/1957Aller 29-407 X 3,034,343 5/1962 Henry et al. 73-116 3,037,348 6/1962Gassmann S0-39.16 3,038,331 6/1962 Lindberg 73--116 3,044,262 7/ 1962Chadwick et al. 60-39.l6 3,077,030 2/ 1963 Carlson 29-407 3,120,0532/1964 Lewis 29-407 WHITMORE A. WILTZ, Primary Examiner.

ROBERT M. WALKER, THOMAS H. EAGER,

Examiners.

UNITED STATES PATENT OFFICE CERTIFICATE 0F CORRECTION Patent No 3 ,252,212 May Z4, 1966 Albert H. Bell III It is hereby certified that errorappears in the above numbered patent requiring correction and that thesaid Letters Patent should read as corrected below.

Column 4 line 6 for The" read [The Column 5, line 17, for "turbines"read turbine line 47, for "matchng" read matching column 6, line 3 for"arrangement" read arrangements column 8 lines 38 to 4l and column 9lines 27 tonSO in the equations for "r each occurrence read r Column 9lines 27 to 30 the left-hand portion of the equation should appear asshown below instead of as in the patent: I

P0 '\|TO column l1 1lines 49 and 50 for "experienec" read experiencecolumn 13 line 44 for "differenttial" read differential Signed andsealed this 7th day of November 1967.

(SEAL) Attest:

EDWARD M.FLETCHER,JR. EDWARD J. BRENNER Attesting Officer Commissionerof Patents

1. A METHOD OF SELECTIVELY MATCHING A GAS TURBINE ENGINE COMPRESSORTURBINE WHEEL TO THE COMPRESSOR TURBINE WHEEL NOZZLE ASSEMBLY,COMPRISING THE STEPS OF CREATING A FIRST PRESSURE DIFFERENTIAL ACROSSSAID NOZZLE ASSEMBLY SUBSTANTIALLY EQUIVALENT TO A PREDETERMINEDPRESSURE DIFFERENTIAL WHICH SAID NOZZLE ASSEMBLY WILL EXPERIENCE WITHINSAID ENGINE DURING SELECTED PERIODS OF NORMAL ENGINE OPERATION,EMPLOYING SAID FIRST PRESSURE DIFFERENTIAL TO FLOW A SUITABLE GASTHROUGH SAID NOZZLE ASSEMBLY, DETERMINING FROM SAID PRESSUREDIFFERENTIAL AND THE VOLUME RATE OF FLOW OF SAID GAS THERETHROUGH THEEFFECTIVE FLOW AREA OF SAID NOZZLE ASSEMBLY, REPEATING THE ABOVE STEPSWITH SUCCEEDING NOZZLE ASSEMBLIES, CLASSIFYING EACH OF THE NOZZLEASSEMBLIES ACCORDING TO THEIR RESPECTIVE EFFECTIVE FLOW AREAS, CREATINGA SECOND PRESSURE DIFFERENTIAL ACROSS SAID TURBINE WHEEL SUBSTANTIALLYEQUIVALENT TO A SECOND PREDETERMINED PRESSURE DIFFERENTIAL WHICH SAIDTURBINE WHEEL WILL EXPERIENCE WITHIN SAID ENGINE DURING SELECTED PERIODSOF NORMAL ENGINE OPERATION, EMPLOYING SAID SECOND PRESSURE DIFFERENTIALTO FLOW A SUITABLE GAS-